1. Field of the Invention
This invention is generally directed to satellite attitude and power acquisition systems and methods, and, more particularly, to satellite attitude and power acquisition systems and methods that are applicable to satellites in a solar wing-stowed configuration.
2. Description of the Related Art
Transporting a spacecraft from the ground to a destination orbit is an integral and crucial part of any spacecraft mission. For example, to insert a spacecraft into a geosynchronous orbit, a launch vehicle typically injects the spacecraft into a low-altitude parking orbit. The spacecraft then performs transfer orbit operations to transfer the spacecraft from the parking orbit to a destination orbit. The transfer orbit is usually performed by firing a liquid apogee motor (LAM) with the spacecraft spinning around a LAM axis to stabilize the spacecraft and to even the thermal and power conditions, or by firing a combination of LAM and xenon-ion propulsion (XIP) thrusters. Once the spacecraft has completed its transfer orbit, it then may enter in-orbit testing and on-station operation.
From cradle to grave, the spacecraft may go through the following phases of operations: separation, transfer orbit operation (including coasting, spin speed change, reorientation and LAM burn), deployment (including antennas, reflectors, solar wings, radiators), acquisition (including power acquisition and attitude acquisition), in-orbit test (including antenna mapping), on-station operation (including normal pointing, momentum dumping, station keeping and station change), and a deorbiting operation.
Typically, spacecraft, such as communication satellites, use multiple separate sets of sensors and control algorithms for different phases of operations. For example, different sets of sensors and/or control algorithms may be used for attitude determination and control for bi-propellant spinning transfer orbit operations versus those that are used for on-station operations. The use of different sensors, attitude determination, and attitude control methods for spinning transfer orbits and on-station operations, respectively, increases the spacecraft weight, sensor and processor complexity, as well as the development cost for spacecraft attitude determination and control systems.
Spinning transfer orbit operations for spacecraft typically may be performed by ground-assisted attitude determination using a spinning earth sensor and a spinning sun sensor set. The measured leading edge and trailing edge of the earth detected by the earth sensor and the measured TOA (time of arrival) of the sun detected by the sun sensor collected and relayed periodically to a ground station. Typically, at least one orbit pass is dedicated to this data collection. A ground orbital operator may then run a ground attitude determination algorithm using these inputs and ephemeris-computed sun and earth positions to determine the spin axis attitude of the spacecraft. This spin axis attitude (the spin phase being still undetermined) is then uploaded to the spacecraft. Next, on-board software may use this spin axis attitude together with the spin phase measured by the spinning sun sensor to complete the 3-axis attitude determination for subsequent spacecraft reorientation or liquid apogee motor (LAM) burn. Power acquisition or sun acquisition are typically performed with a wide-field-of-view sun sensor, an orthogonal set of narrow-field-of-view slit sun sensor, or an orthogonal set of spinning slit sun sensor.
On-station spacecraft operations typically use different sensors, such as a staring earth sensor assembly (STESA) and a wide field of view (WFOV) sun sensor assembly (SSA), and/or a star tracker for attitude determination; and WFOV sun sensor for power acquisition or sun acquisition. Thus, the sensors used for transfer orbit operations may lie dormant for the entire time that the spacecraft is on station. The number of sensor types used and the number of sensors used, increase the hardware and development cost, increase weight and launch cost, and complicate the mission operation. In addition, some spacecraft have configurations and equipment that may make it difficult in some situations to provide a clear field of view for some sensors, such as, for example, a WFOV SSA, which spans a diamond of about 120×120 deg.
The present invention is directed to overcoming one or more of the problems or disadvantages associated with the prior art.